The present invention relates to a gas turbine combustion chamber, especially for an aircraft gas turbine or for a stationary gas turbine. As is known from the state of the art, air supplied by the compressor is used for cooling the combustion chamber wall. Without cooling, the compressor chamber wall would reach temperatures up to 2,600 K. Obviously, no materials are available which are capable of resisting such temperatures without cooling.
Disadvantageously, a relatively large part of the combustion air (between 30 and 50 percent of the total combustion air) must be used for cooling the combustion chamber.
With highly advanced, single-annular combustion chambers provided with lean premix burners, it is currently assumed that 60 to 70 percent of the combustion air (W30) is available for premixing in the lean premix burners to reduce NOx air pollutant emissions.
In the future, turbofan engines will possess significantly higher thermal efficiency and better propulsive efficiency than the state of the art.
As a result of both measures, the core engine must be operated with higher combustion chamber temperatures (entry and exit) and with a lower total air-fuel ratio, i.e. using a richer mixture.
In future combustion chambers, with a given constant cooling air efficiency, less air will, therefore, be available for premixing and the NOx air pollutant emissions will inevitably increase as a result of the higher combustion temperatures.
Accordingly, the only possibility with conventional engines is to provide sufficient combustion air to enable the premix to be leaned and the combustion temperatures correspondingly reduced. Here, efficiency of the air-cooling process can be increased to a limited extent only. It is not expected that the cooling air share can be decreased clearly below 30 percent.